A solar array or array, as defined herein, pertains to a structure that is attached to certain spacecraft vehicles or satellites, to provide power for spacecraft operations. The solar array is stowable into a small volume for shipment and launch, and is deployable when in space to expose a large surface area of photovoltaic (PV) solar cells to the sun, to collect solar radiation and convert it into the electrical power necessary to operate the spacecraft.
Power systems for space applications face numerous design constraints including criteria to minimize weight, minimize stowed volume, maximize beginning or life to end of life performance, and minimize cost. In certain prior art applications of solar arrays, the structure consists of flat rigid honeycomb panel substrates to which the solar cells are mounted that are configured for stowage by means such as hinges which will permit the panels to be folded against each other to minimize the dimensions of the array in the stowed configuration. The solar array typically comprises one or more solar panels electrically and mechanically attached to each other and to the spacecraft. Each solar panel in an array typically comprises numerous individual solar cells, which are usually laid out in rows and connected together electrically at their adjacent edges to form a two-dimensional array. The mechanical components and large stiff panels associated with rigid panel arrays involve added weight which is desirable to minimize. An example of such an array is shown in: Everman et al U.S. Pat. No. 5,487,791.
In order to allow for further reduction in the deployable solar arrays weight and stowed volume, the solar cells can be mounted to a light weight flexible substrate, or blanket instead of the large and heavy rigid honeycomb panels. Various flexible solar cell blanket substrates have been used, such as those fabricated from a fiberglass mesh or thin polymeric sheet upon which are bonded the numerous crystalline solar cells. Flexible photovoltaic (PV) blanket solar arrays are typically limited to crystalline solar cells packaged onto a long continuous roll or pleated and folded stack that is attached to and deployed by a separate deployment boom actuator, hub structure or other deployable structure requiring external motor power or material strain energy for deployment motive force.
The number of solar cells that must be employed on a solar array is a function of the anticipated spacecraft power demand and the efficiency of the cells. High-efficiency solar cells are typically employed to reduce the area of photovoltaics required by a specific spacecraft. This reduces panel (or flexible PV blanket) area and thus overall mass from the required supporting structure and minimizes the volume of the stowed power system. But such cell devices are extremely expensive, and in many cases cost impactive for certain applications. Solar cells are by far the most expensive component of a solar array. Because system cost and mass both increase directly with the number of solar cells employed, there is considerable economic incentive to reduce the quantity of solar cells that a spacecraft must carry on an array.
To reduce solar array cost and more mass efficiently shield the solar array from radiation exposure, reflective or refractive concentrator elements may be used to reduce the number of cells. Lightweight reflective surfaces have been used in various combinations with known rigid solar panels to produce power with fewer solar cells. Lightweight refractive optics develop by Fraas and O'Neill, U.S. Pat. No. 5,344,497 and U.S. Pat. No. 5,505,789, including using lenses such as point-focus or line-focus Fresnel optics to refract the solar illumination onto the cells, have also been used in various combinations with known rigid solar panels to produce power with fewer solar cells. By using relatively inexpensive Fresnel lens optics to collect the sunlight and to focus it onto much smaller solar cells, the cost and weight of the cells for an equivalently powered solar array are dramatically reduced. By using very high efficiency cells, the required array area is minimized, reducing overall system weight and launch volume.
Despite the many advantages of the Fresnel lens concentrating solar array previously invented by O'Neill and Fraas, this array still has shortcomings in need of improvement. Specifically, the Fresnel lens is presently made from a space-qualified, optically clear silicone rubber material (e.g., Dow Corning DC 93-500). In the late 1990's, NASA's New Millennium Deep Space 1 spacecraft implemented a line-focus Fresnel lens refractive concentrator rigid panel solar array. For this concentrator array, the 250-micron thick/thin rubber lens was laminated to a thin 80-microns thick ceria-doped glass superstrate to maintain the required arch shape of the lens assembly. The glass superstrate was required to provide structural strength and stiffness properties in the stowed and deployed configurations, but was not required for the optical functioning of the lens. Unfortunately, the implementation of the glass superstrate increases the weight, cost, launch volume, and fragility of the lens assembly. If the glass superstrate was not used for the Deep Space 1 concentrator solar array then the lens would not maintain its shape, even in the zero-gravity environment of space. The glass/silicone Fresnel lens used on Deep Space 1 also required a supporting structure to properly position the lens above the solar cells. This lens support structure added further weight, cost, and complexity to the solar power system. The glass/silicone Fresnel lens construction used on Deep Space 1 is also not flexible enough to be flattened for compact launch stowage, resulting in a higher than desired launch stowage volume. Finally, the glass/silicone Fresnel lens is affected by the difference in thermal expansion coefficients of the glass and silicone layers, causing either stresses or strains in the lens during temperature variations which occur when the satellite moves in and out of the Earth's shadow.
One means of addressing some of the problems associated with the glass superstrate in the presently used glass/silicone Fresnel lens is to make the polymer lens from a stronger, thicker material, obviating the structural need for the glass arch. Many stronger, thicker polymer lens materials that are different to DC-93500, such as Telfon and other flouro-polymers, have been evaluated under NASA and DoD programs, but have had displayed limited survivability after exposure to the combined space environment. In U.S. Pat. No. 5,496,414, Harvey et al. describes one means of stowing and deploying such a monolithic polymer lens. In U.S. Pat. No. 5,578,139, Jones et al. describes another means of stowing and deploying such a monolithic polymer lens. However, these prior art lenses must be thick enough and strong enough to be self supporting during ground testing, and therein lies yet another disadvantage. The lens thickness required to be self-supporting under one gravity acceleration ground testing is typically 250 microns or more for an 8 cm lens aperture width. Since the density of fluoro-polymers, and other possible alternative lens materials, is about double the density of the normal DC93-500 silicone rubber lens material, and the total lens thickness is about the same, the flouro-polymer lenses weigh about twice as much as the silicone lenses. Thus, even with the added weight of the glass arch superstrate, the Deep Space 1 glass/silicone lens construction is typically lighter than a potential monolithic fluoro-polymer lens. Most importantly and as previously stated, the monolithic fluoro-polymer lens material need for the Harvey and Jones design embodiments does not have the proven successful space flight history and heritage of the silicone lens material.
U.S. Pat. No. 6,075,200, O'Neill, describes a single monolithic DC93-500 silicone material Fresnel lens that is stretched and strained to provide strength and stiffness in the deployed configuration. The O'Neill invention uses one-dimensional lengthwise tension to support the thin lens material in the space environment and produce a precise lens optical shape while retaining lightweight and the use of space proven DC93500 lens materials. This tensioning approach enables the lens to maintain an ideal, arched, curvilinear shape, with absolutely no aperture blockage over the full stretched length of the lens. The O'Neill invention employs an arched shaped structure and spring powered hinge mechanisms at each end of each relatively short lens segment. Each lens segment, with integral arch structures and spring hinge mechanism at each lens end allows the lens to be folded flat against the photovoltaic receiver/waste heat radiator assembly, for minimal launch volume. Once on orbit in space, the spring powered mechanisms on each lens arch structure on each end of each lens segment deploy and pop-up into place, thereby lightly tensioning the lens in one direction. By maintaining a small lengthwise stretching force on the lens, either with springs or flexible structure, small differential thermal expansion and contraction of the lens relative to the solar cell receiver/radiator structure is accommodated. A major drawback of the O'Neill embodiment of U.S. Pat. No. 6,075,200 is that its only particularly suited for integration of strength lens segment assemblies onto conventional rigid honeycomb panels. The tensioning loads of the lens results in a compressive load onto the base solar cell receiver/radiator structure, whereby a more rigid and heavier receiver/radiator structure of honeycomb construction is required. The requirement for a more rigid and heavier, or honeycomb, structure prohibits the ability to accordion fold or roll the entire solar array panel assembly and create an extremely compact stowage volume, a feature particularly desired by the end-user. Additionally, the O'Neill embodiment requires a number of parts and complex mechanization to achieve stowage and deployed states which further increase mass and cost. Finally, the O'Neill embodiment involves bi-direction folding of discrete lens elements and arches towards each other. Bi-directional folding does not lend itself towards rolled stowed packaging architectures and does not produce as compact of stowed package. The proposed embodiment contained herein involves unidirectional folding/collapsing of the lens segments and supporting arches which lends itself to both compact rolled and accordion folded flat-pack stowage architectures, and overall produces a more compact stowage.
The O'Neill U.S. Pat. No. 6,075,200 fails to adequately address another critical problem with stretched lens concentrator embodiment, which being the differential thermal expansion between the lens material and the support structure, especially in the direction of greatest linear dimension. In the earth orbital space environment extreme variations in temperature are realized as the deployed solar array enters and leaves the shadow of the earth. In the earth's shadow temperatures as low as −180 degrees Celsius are endured, while out of the earth's shadow and exposed directly to the solar illumination temperatures as high as 110 degrees Celsius are endured. In earth orbital space applications the optics need to operate reliably and maintain positional accuracy while out of the earth's shadow and exposed directly to the solar illumination. Operating temperatures beyond earth orbital applications, and near the solar system outer planets (at Jupiter or Saturn at distances of 5 to 7 astronomical units, respectively), can be as low as −180 degrees Celsius while exposed directly to the solar illumination. For these outer planets missions it is imperative the refractive lens concentrator optics have the ability to accommodate broad temperature extremes while maintaining precise positional alignment of the optics and with negligible distortion. The presently used silicone lens material expands and contracts at an enormous rate with temperature (more than 300 parts per million per degree Centigrade) with respect to its supporting structure and tensioning system. An unrestrained 30 cm (1 foot) long lens will expand and contract more than 2 cm (8%) in length during an outer planets mission temperature excursion from −180 to 110 degrees Celsius. In contrast, a typical graphite/epoxy space structure will expand and contract several hundred times more slowly with temperature than the lens material. The differential thermal expansion problem is somewhat addressed in the O'Neill embodiment, but only for earth orbital applications were very small temperature extremes are realized. However, the differential thermal expansion problem is significant for an outer planets mission where very large operating temperature extremes are apparent, and as such must be addressed for an acceptable stretched lens space solar concentrator, but the prior art does not teach a solution to this problem.
Accordingly, several objects and advantages of the proposed embodiment are to provide improved refractive lens concentrator solar arrays for space power applications, said improved solar array concentrators providing lower mass, more compact stowage volume for launch, lower cost, more reliable deployment on orbit, flexible stowage packaging of either compact rolled or accordion folded compactly stowed architectures, and sustainable and reliable operation under very broad temperature exposure.